Gas turbine engines, such as those used on jet aircraft, generally comprise an air intake port, a fan mounted on a hub near the air intake port and surrounded by a fan case which is mounted within an engine housing or nacelle, a low pressure compressor (LPC) section, an intermediate section aft of the LPC section, a high pressure compressor (HPC) section, a combustion chamber or combustor, high and low pressure turbines that provide rotational power to the compressor blades and fan respectively, and an exhaust outlet. The HPC section includes a number of stages of circumferentially arrayed, stationary stators and rotating blades. Each stator extends radially outward between a rotary disk stack and the inboard surface of an inner case. The inner case may be circumferentially segmented into multiple inner case segments joined by inner case flanges. These inner case flanges are subjected to significant stress during operation.
An outer case and the inner case define the outer and inner circumferentially arranged walls of an air bleed manifold surrounding at least a portion of the HPC. The air bleed manifold defines a cavity and is supplied with air from two sources: cold air from the HPC sixth stage bleed ducts and hot air from the diffuser strut flow from the aft hub.
The cold and hot air mix in the air bleed manifold cavity. The hot air creates maximum temperature and lifting problems due to high temperature gradients and hot air mixing, which raises the steady state temperature of the inner case flanges, even when a heat shield is used. As a result, the hot air can reduce engine part life, including the life of the inner case flanges. The present disclosure is directed at solving this problem.